The aspects of the disclosed embodiments concern an aircraft fuselage section made of composite material having an approximately constant internal profile along the longitudinal axis of the aircraft. The disclosed embodiments also concern a manufacturing process for a skin made of composite material for such a fuselage section.
The disclosed embodiments have applications in the field of aeronautics, particularly the field of constructing fuselage sections made of composite material.
Traditional aircraft fuselages are metal. A metal fuselage is comprised of metal panels mounted and attached around an internal structure, which is also metal, called the internal skeleton of the aircraft. Once assembled, these metal panels form the skin of the aircraft. Such metal fuselages have the disadvantage of being heavy, since the internal structure and the skin are metal.
To reduce the weight of the fuselage, aeronautics manufacturers have replaced certain metal elements with elements made of composite material. Composite materials are mainly used to make the skin of the aircraft fuselage. This skin of composite material is made from strips of dry fibers or fibers impregnated with a thermosetting resin. For example, strips of pre-impregnated fibers are placed on/in a mold to be formed, then heated with the mold. The heat causes the resin to polymerize, which allows the fiber reinforcement to take the shape of the mold. After it cools, the mold is removed.
Whatever the type of fuselage, metal or composite material, an aircraft is subject to aerodynamic forces in flight. The aerodynamic forces on the wing section of the aircraft cause flexure of the fuselage. This flexure is proportional to the proximity of the wing root section. The farther the fuselage area in question is from the wing root section, the less flexure. The flexural torque is even close to zero in the nose of the aircraft. On the other hand, in the tail of the aircraft, the flexural torque is never zero due to the presence of the horizontal plane, which generates aerodynamic forces.
An example of an aircraft with the traditional shape is shown in FIG. 1, with the profile of the flexural torque for the aircraft when it is in flight superimposed. This figure shows that the maximum flexural torque My is located at the root of the wing section.
Logically, there is a variation in the thickness of the fuselage skin that corresponds to the profile of the flexural torque My of the fuselage. Thus, the thickness of the skin increases from the nose of the aircraft up to the wing root section, and then decreases from the wing root section to the tail of the aircraft.
FIG. 2 shows the aircraft in FIG. 1 with the different forces corresponding to this flexural torque My exerted on the fuselage. More specifically, the flexural torque My causes tractive stress FT on the top 10 of the fuselage, i.e., on the top panels of the fuselage. It causes compressive stress FC on the bottom 11 of the fuselage; the bottom includes all of the lower panels of the fuselage, opposite the top and facing the ground. The torque My causes shearing stress T, or torsional moment on the side panels 12 of the fuselage.
To ensure better resistance against these stresses, the skin of the aircraft is reinforced at the panels most stressed. The thickness of the aircraft skin is therefore irregular.
It is also known that a stiff structure makes it possible for more tractive forces than compressive forces to pass for the following reasons:                the characteristics of the composite materials used to make the fuselage are better in traction and        once the skin has shrunk due to the compressive stress, the working width is reduced.        
Consequently, to take this characteristic into account, the thickness of the fuselage covering, i.e., the skin, gradually increases from the top 10 to the bottom 11. For example, the skin on the top goes from 1.4 to 1.6 mm in 10 interframes (one interframe equals 533 mm) and the skin on the bottom increases from 2 to 2.8 mm in 20 interframes (an interframe is still equal to 533 mm). An example of the skin thickness is shown in FIG. 3. FIG. 3 shows a curve C10 for the thickness of the skin on the top and a curve C11 for the thickness of the skin on the bottom. The change in skin thickness from top to bottom is gradual and follows the profile of the flexural torque My.
On a traditional, i.e., metal aircraft, the changes in skin thickness are made by machining pockets. For this, sheet metal with the maximum thickness is used, i.e., its thickness corresponds to the maximum thickness of the skin on the bottom. Then, more or less deep pockets are machined, chemically or mechanically, depending on the skin thickness desired.
Such a technique makes it possible to obtain fuselage panels of variable thickness based on the location of said panels on the fuselage. With this technique, the variation in thickness is on the internal profile of the fuselage. For aerodynamic reasons, the sheet metal designed to form the fuselage panels is machined on its inside, that is, on the side of the panels located in the interior of the fuselage. It can therefore be seen that the internal profile of the fuselage is not constant; it can vary in proportion to the variation in the thickness of the skin.
Now, this variation in the internal profile of the skin causes problems attaching the frames and stringers forming the internal structure of the fuselage. It will be recalled that, in an aircraft, the structural elements that form the skeleton, or internal structure, of the aircraft require the use of additional pieces to improve the mechanical hold of these structural elements. These additional pieces can be stiffeners. Stiffeners are profiled pieces attached to the structural elements of the aircraft, for example, panels, to transfer loads or stabilize elements. Stiffeners can be frames, which are radial structural elements of the fuselage or stringers, which are longitudinal structural elements of said fuselage. The frames and the stringers are used particularly to stiffen the skin and stiffen certain specific areas of the aircraft, such as the door frames. The stiffeners, frames or stringers, can have different shapes: for example, a Z, T, J or Omega shape. To offset the variation in the internal profile of the fuselage, shims or wedges are placed under the stringers or frames at the places where the pockets are the largest, and thus where the skin is least thick, to allow stringers and frames of the same size to be attached.
FIG. 11 shows an example of a fuselage panel 13 to which are attached stringers 14 and frames 15. In this example, the stringers have an Omega shape, and the frames are Z-shaped attached with clips 16.
In the case of a fuselage made of composite material, there is no wedge under the stringer or frame. In fact, the draping of the strip of fibers is too complex and too expensive to allow wedges just at the sites of the frames and stringers. It would be possible to consider putting wedges under the stringers and frames by a process of locally superimposing strips of fibers called “the pad-up process,” but the mechanical performance of such superimposing would be very poor, because it would be done without interlacing the plies.
In addition, the rules for ply drop-off of composite material are difficult, or even impossible, to apply to the intersections between stringers and frames.
In composite material technology, variations in skin thickness can be made by a process of ply drop-off. In this process, the variations in thickness are obtained by varying locally the number of strips of fibers superimposed by interleaving and offsetting said strips of fibers. As shown in FIG. 4, strips of fibers of different sizes are interleaved between strips of fibers of maximum sizes to create interlaced plies. These plies make it possible to vary the thickness of the skin.
However, this ply drop-off process is slow and difficult (many ply stops and starts). What is more, with such a process, the cost of manufacturing the fuselage is directly connected to the number of variations in thickness. In effect, if a ply drop-off is being made in the longitudinal direction of the aircraft, i.e., a jump in thickness, on the stringers, the stringers must be plunged or sloped so each stringer can be attached according to the profile of the jump in thickness. Then the stringers cannot be produced by pultrusion. Remember that the pultrusion manufacturing process is a process of manufacturing stringers continually, with the stringers being cut later to the desired length. Such a process has the advantage of being inexpensive, since all the stringers produced are identical; however, this process requires that the stringers have a constant cross section. If the inside of the skin has ply drop-offs the dimension of the stringers must necessarily be variable. Therefore, pultrusion cannot be used to produce stringers, which makes the manufacturing cost expensive.
If there is a ply drop-off in the longitudinal direction on the frame of the aircraft, the frames must be modified. But the size of an aircraft cabin is predetermined; the total height of the skin and the frame is therefore limited. Consequently, increasing the skin thickness necessarily means modifying the size of the frame. In particular, in the case of clipped frames (shown in FIG. 11), the geometry of the clip must be modified by reducing its height to offset the increase in skin thickness. Such a modification in the geometry of the clip has the disadvantage of not being suited for composite structures, where it is generally preferred that the attachment pieces be built in.
Indeed, on a composite structure, we are attempting to reduce to the maximum the number of attachments to be placed in the structure, since drilling is delicate, requiring special drills with very expensive diamond coatings and entailing a risk of delamination when the drill comes out and bad onboard behavior of holes for lack of plasticizing.
In a structure made of composite material, we are trying to use to the maximum integrated frames for which the clips are integrated into the frame preform. This avoids attachments between clips and frames, which gives better mechanical hold, saves weight and consequently assembly cost. These integrated frames can be F-shaped frames, as shown in FIG. 5. These F-shaped frames are generally manufactured according to RTM technology, which requires an expensive mold. Consequently, each jump in thickness on the inside of the skin entails the manufacture of a frame with different dimensions and, hence the design of a special mold to make that frame located at the jump in thickness. It follows that this technique, while powerful, requires substantial tooling costs.
Another solution could consist of designing a mold for a frame with a smaller diameter and placing a shim between the foot of the frame and the skin, when the skin is thinner. However, this solution would entail the cost of manufacturing a specific piece to brace the frame; this bracing, for a structure made of composite material, is liquid bracing that requires polymerization, which is also expensive. Such a solution would also reduce the mechanical performance of the joint, since it adds an offset of neutral fibers.
FIG. 6 shows different examples of frames for the different technologies explained above. These frames 20 are mounted on a skin 2, more or less thick, based on the examples. In particular, part A of FIG. 7 shows a frame 20a with a C-shaped profile, clipped to a thin skin 2a by means of a clip 22. Part B shows a built-in frame 20b, attached to a thin skin 2a. Part C shows a built-in frame 20c with a shim 23 between the thin skin 2a and said frame; the difference in height between this frame 20c and frame 20b is offset by the shim 23. Part D shows the integrated frame 20c attached to a thick skin 2b. These four types of frames correspond to the different embodiments described earlier.
It will be understood from the preceding that none of the current techniques makes it possible to optimize the thickness of a skin made of composite material in the longitudinal direction to absorb the flexure of the fuselage, without significantly increasing the cost of manufacturing the fuselage (either by increasing the cost of the stringers and frames, or by increasing the cost of the skin).